This invention relates generally to gas turbine engines and, more particularly, to methods and apparatuses for reducing thrust bearing net axial load caused by pressurized rotor-stator cavities.
Gas turbine engine rotor blades typically include a compressor, a combustor, and a high-pressure turbine. In operation, air flows through the compressor and the compressed air is delivered to the combustor wherein the compressed air is mixed with fuel and ignited. The heated airflow is then channeled through the high-pressure turbine to facilitate driving the compressor.
More specifically, High Pressure Centrifugal Compressors (HPCC) may contribute a relatively high percentage of the axial load to a high-pressure rotor thrust bearing. Gas turbine engines with multiple stage compressors and turbines may somewhat result in a balance axial load. However, utilizing a single stage HPCC, wherein the HPCC radius is larger than the turbine radius, may result in a relatively high forward axial load on the turbine.
To facilitate reducing the axial load on the turbine, at least one known gas turbine engine attempts to create a high flow and free vortex environment. However, this is relatively difficult to achieve in a typical HPCC rotor-stator cavity due to stator drag and limitations on flow for operability. As a result, pressure decreases may be realized at primarily at low radii with high flowing systems resulting in relatively small changes in load.